Rocket engine with integrated oxidizer catalyst in manifold and injector assembly

ABSTRACT

A rocket engine has: a combustion chamber having a chamber inlet for receiving an oxidizer and a chamber outlet for expelling combustion gases in an environment outside the combustion chamber; a manifold having a manifold inlet fluidly connectable to a source of the oxidizer and a manifold outlet; a catalyst having a catalyst inlet fluidly connected to the manifold outlet and a catalyst outlet; and an injector plate having a injector inlet fluidly connected to the catalyst outlet and an injector outlet fluidly connected to the chamber inlet.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims priority from U.S. provisional application No.62/798,679 filed Jan. 30, 2019, the entire contents of which areincorporated by reference herein.

TECHNICAL FIELD

The application relates generally to rocket engines and, moreparticularly, to hybrid propellant rocket engine configurations.

BACKGROUND OF THE ART

The development of new high-performance fuel/oxidizer combinations forrocket propulsion applications must take into consideration severalimportant parameters such as the propellants specific impulse,storability, ignition delay, stability and toxicity. Hybrid rocketpropellants have the potential to address some of the drawbacks ofliquid alternatives, mainly because of their relative simplicity, lowdevelopment cost, the ability to control their thrust and re-ignite themfollowing partial burns, and because it may alleviate the requirement tomatch the momenta of the dual propellant streams during throttlingoperation.

Standard hybrid rocket engines have not yet found commercial spaceflightapplications because they suffer from slow solid-fuel regression rates,low volumetric loading and relatively poor combustion efficiency,compared to existing alternatives. This may be due to their inability toburn at a constant oxidizer to fuel ratio, which may result in pooraverage combustion performance. This oxidizer to fuel ratio shiftthroughout the duration of the burn, meaning that the rocket engine'sspecific impulse cannot be maintained at its peak value, since the fuelis not burning at its optimal stoichiometric ratio.

There are currently four common ways of powering injection of propellantinto a combustion chamber of a rocket engine. These four injection waysare divided into two main categories. The first category includespressure-fed power cycle, which demonstrated the most potential forhybrid rocket propulsion systems, as it is the one with the leastcomponent complexity and moving parts. The second category, whichencompasses the other three cycles (staged combustion, expanders andgas-generator cycles) and their derivatives, all rely on turbo-pumps toinject propellants into the combustion chamber. Turbo-pumps increasecomplexity since they require gas-generators which incorporate their owninjector set, regulators and so on.

Propulsion systems using pressure-fed cycles have practical limits onpropellant pressure, which in turn limits combustion chamber pressure,thus minimizing performance. High pressure propellant tanks requirethicker walls and stronger alloys which makes the vehicle tanks heavier,reducing performance and payload capacity.

SUMMARY

The disclosed thrust chamber assembly combined with the disclosed highvolumetric specific impulse solid polymer fuel may enable constantoxidizer to fuel ratio, maintain high combustion performance, and highideal characteristic velocity. The solid polymer fuel within the thrustchamber assembly incorporates means and methods to control its dropletsvaporization mechanism through a change of its thermorheologicalproperties, which may result in a constant fuel mass flow release. Thosethermorheological properties may be tailored to ensure a nearstoichiometric combustion performance through the duration of the rocketengine's operations, which may yield an overall high specific impulsewith enough margins to enable orbital flight.

The combustion process of liquefying solid polymer fuels may involvesequential vaporization and combustion of droplets vapors in the gasphase. The rate at which liquid fuel droplets vaporize and combust isdependent on thermorheological properties of the fuel, most notably thedroplet sizes and the rates at which heat can be transferred to theliquid fuel droplet surfaces and the mixing characteristics in thecombustion chamber. Layer by layer, the solid fuel is vaporized, and theliquid fuel droplets are then consumed. Even for smaller droplets, thephenomena are the same, but at a much faster rate with good mixing.

The disclosed high volumetric specific impulse solid polymer fuels maydrastically reduce the mechanical complexity of the propulsion systems.Additionally, the complexity of the system is largely determined by thetype of power cycle used for the rocket propulsion system.

The high volumetric specific impulse of the fuel may allow for a majorvolume reduction of the oxidizer tank and the combustion chamber thusincreasing the vehicle' structural margins, which may increase stagemass ratio. This increase of stage mass ratio may compensate for themass losses that may be required for pressure-fed pressuring gascontained within a tank. The disclosed propellants demonstratedhypergolic characteristics with various oxidizers may give missionplanners more flexibility with mission flight paths designs and givesvehicle designers more options for future vehicle developments withenhanced capabilities.

The use of fuels with high volumetric specific impulse may allow a majorvolume reduction of the oxidizer tank and the combustion chamber thusincreasing the vehicle's structural margins, increasing the launchvehicle's payload mass. This may allow for a reduction of structure massdue to lower pressure needs and for a reduction of propellant mass dueto higher volumetric specific impulse.

The present disclosure provides a thrust chamber assembly incorporatinga solid polymer fuel having a mass flow that may be unaffected by thevariation of the oxidizer's flux for the full duration of the rocketengine burn, as the fuel's thermorheological properties may becontrolled throughout a length of the latter as a consequence ofdifferent processes that can involve microcrystalline structure,chemical additives, and others, or a combination of the said processes.As a result, the high volumetric specific impulse solid polymer fuelsmay drastically reduce the mechanical complexity of the propulsionsystems, which is an inherent problem for any orbital vehicle.

The disclosed rocket engine might drastically reduce the mechanicalcomplexity of propulsion systems, which is an inherent problem for anyorbital vehicle. The mechanical complexity is largely determined by thetype of power cycle used for the rocket propulsion system.

In one aspect, there is provided a rocket engine comprising: a housinghaving a longitudinal axis and defining a combustion chamber, thehousing defining an inlet and an outlet of the combustion chamber, aflow passage extending from the inlet to the outlet within the housing,the inlet fluidly connectable to a source of oxidizer, the outletopening to an environment outside the combustion chamber for expellingcombustion gases, a first fuel having of a first solid propellant and asecond fuel having of a second solid propellant, the first and secondfuels located within the combustion chamber, the first solid propellanthaving a regression rate greater than that of the second solidpropellant.

The first and second fuels are axially offset from odne another relativeto the longitudinal axis. Each of the first and second fuels may includeat least one annular disk. The second fuel may include three annulardisks of the second solid propellant and the first fuel may include twoannular disks of the first solid propellant, each of the two annulardisks of the first solid propellant may be sandwiched between two of thethree annular disks of the second solid propellant. The first and secondfuels may be radially offset from one another relative to a longitudinalaxis of the housing. The first fuel may be located radially inwardly tothe second fuel relative to the longitudinal axis. The first fuel may bea tube. The inlet may include at least one aperture defined through thehousing, the at least one aperture may be located axially between thefirst and second fuels and the outlet relative to a longitudinal axis ofthe housing. The first and second fuels may be located axially betweenthe inlet and the outlet of the combustion chamber. A third fuel may belocated within the combustion chamber, the third fuel may have a thirdsolid propellant having a regression rate different that those of thefirst and second solid propellants.

In another aspect, there is provided a method of operating a rocketengine, comprising: receiving an oxidizer within a combustion chamber;exposing first and second fuels to the received oxidizer, the first fuelhaving a first solid propellant, the second fuel having a second solidpropellant having a higher viscosity than that of the first solidpropellant; and expelling combustion gasses created by a reactionbetween the received oxidizer and the first and second fuels.

Receiving the oxidizer may include receiving the oxidizer along an axialdirection relative to a longitudinal axis of the rocket engine.Receiving the oxidizer may include receiving the oxidizer via an inlet.Expelling the combustion gasses may include expelling the combustiongasses via an outlet, the first and second fuels may be located axiallybetween the inlet and the outlet relative to a longitudinal axis of therocket engine. Receiving the oxidizer may include receiving the oxidizerin a direction having a radial component relative to a longitudinal axisof the rocket engine. Receiving the oxidizer in the direction having theradial component may further comprise receiving the oxidiser in thedirection further having a circumferential component relative to thelongitudinal axis.

In another aspect, there is provided a rocket engine having a combustionchamber with at least two kinds of solid fuels that differ by theirrheological properties. This may allow to yield an optimal oxidizer tofuel ratio during the total duration of the burn and may thus allow toachieve ideal characteristic velocity. Having two fuels with differentviscosities may offer the ability to alter the shape of the fuelstacking during the combustion phase to obtain some coolingcharacteristics. The disclosed rocket engine has a thrust chamberassembly, also referred to as the combustion chamber, incorporating astacking of several solid fuels with the fuel's mass flow that may beunaffected by the variation of the oxidizer's flux throughout theduration of the burn, as the fuel's viscosity may be varied between thedifferent stacking.

Effectively guiding a rocket launch vehicle to its intended orbitrequires precise guidance and control and the control of the directionof exhaust from the vehicle's rocket engine (referred to as ThrustVector Control or TVC) is important to achieving this control. Onemethod to control the direction of exhaust gases is through the preciseinjection of a fluid (gas or liquid) into the expansion section of therocket nozzle.

To achieve this thrust vectoring effect with higher reliability, and toenable assembly and manufacturing costs of individual engines, there isdisclosed herein a concept for embedded injection ports which mayintegrate directly into the sides of the expansion section of the rocketnozzles. The part may be manufactured using fused metal depositionmanufacturing and will have an injector apertures to inject a workingfluid into the exhaust gases, a port to accept a standardized valve, anda shape allowing it to be easily integrated into the side of the nozzle.Since the part may be add 3D printed, it may match the complex curvatureof the nozzle and holes drilled in the nozzle to accept the part. Aswell, the part may use features which may facilitate bonding andlaminating to the ablative liner of the surface, which may allow for abetter and stronger seal to be made with the nozzle extension. Theintegrated and monolithic nature of the part may allow for the thrustvector control plumbing to be quickly assembled with the engine.

In one aspect, there is provided a rocket engine comprising: acombustion chamber having a chamber inlet for receiving an oxidizer anda chamber outlet for expelling combustion gases in an environmentoutside the combustion chamber; a manifold having a manifold inletfluidly connectable to a source of the oxidizer and a manifold outlet; acatalyst having a catalyst inlet fluidly connected to the manifoldoutlet and a catalyst outlet; and an injector plate having a injectorinlet fluidly connected to the catalyst outlet and an injector outletfluidly connected to the chamber inlet.

In another aspect, there is provided an integrated injector assembly forinjecting an oxidizer within a combustion chamber of a rocket engine,comprising an inner wall defining an inlet fluidly connectable to asource of an oxidizer, an outer wall spaced apart from the inner wall todefine a cavity therebetween, the outer wall defining an outlet forinjecting a catalyzed oxidizer within the combustion chamber, and acatalyst located within the cavity.

In yet another aspect, there is provided a method of supplying anoxidizer within a combustion chamber of a rocket engine, comprising:receiving an oxidizer; distributing the received oxidizer whilecatalyzing the received oxidizer; and injecting the catalyzed oxidizerwithin the combustion chamber.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic view of a rocket engine system in accordance withone embodiment;

FIGS. 1a to 1c are schematic cross-sectional views of a rocket engine inaccordance with one embodiment taken at different time steps (1a: t=0;1b: t=t1; 1c: t=t1+Δt);

FIGS. 2a to 2c are schematic cross-sectional views of a rocket engine inaccordance with another embodiment taken at different time steps (2a:t=0; 2b: t=t1; 2c: t=t1+Δt);

FIGS. 3a to 3c are schematic cross-sectional views of a rocket engine inaccordance with another embodiment taken at different time steps (3a:t=0; 3b: t=t1; 3c: t=t1+Δt);

FIGS. 4a to 4c are schematic cross-sectional views of a rocket engine inaccordance with another embodiment taken at different time steps (4a:t=0; 4b: t=t1; 4c: t=t1+Δt);

FIGS. 5a to 5c are schematic cross-sectional views of a rocket engine inaccordance with another embodiment taken at different time steps (5a:t=0; 5b: t=t1; 5c: t=t1+Δt);

FIGS. 6a to 6c are schematic cross-sectional views of a rocket engine inaccordance with another embodiment taken at different time steps (6a:t=0; 6b: t=t1; 6c: t=t1+Δt)

FIG. 7 is a schematic cross-sectional view of a rocket engine inaccordance with another embodiment;

FIG. 8 is a schematic top partially transparent view of an injectorplate of the rocket engine of FIG. 7;

FIG. 9 is a schematic fragmented cross-sectional view, taken on a planecontaining a longitudinal axis of the rocket engine of FIG. 7, andillustrating an injector plate in accordance with another embodimentthat may be used with the rocket engine of FIG. 7;

FIG. 10 is a schematic cross-sectional view of a rocket engine inaccordance with another embodiment

FIG. 11 is a schematic fragmented cross-sectional view, taken on a planecontaining a longitudinal axis of the rocket engine of FIG. 10, andillustrating the injector plate used with the rocket engine of FIG. 10;

FIG. 12 is a schematic top partially transparent view of an injectorplate in accordance with another embodiment that may be used with therocket engine of FIG. 10;

FIG. 13 is a schematic cross-sectional view of a rocket engine inaccordance with another embodiment;

FIG. 14 is a schematic top partially transparent view of an injectorplate of the rocket engine of FIG. 13;

FIG. 15 is a schematic three dimensional view of a portion of a rocketengine in accordance with another embodiment and incorporating anintegrated injector assembly;

FIG. 16 is a schematic cross-sectional view of a rocket engine inaccordance with another embodiment;

FIG. 17 is a schematic cross-sectional view of a rocket engine inaccordance with another embodiment; and

FIG. 18 is a schematic cross-sectional view of a nozzle in accordancewith one embodiment that may be used with any of the rocket engines ofFIGS. 1-17; and

FIG. 19 is a schematic view of a control system for a rocket enginesystem.

DETAILED DESCRIPTION

Referring to FIG. 1, a hybrid rocket engine system is shown generally at100. A hybrid rocket engine combines a solid propellant as fuel and aliquid oxidizer, hence the “hybrid” terminology. The rocket enginesystem 100 includes a rocket engine 102 defining a combustion chamber104 and a convergent-divergent nozzle 106 fluidly connected with thecombustion chamber 104 and in which combustion gases generated in thecombustion chamber 104 may be accelerated from a subsonic speed to asupersonic speed.

The nozzle 106 may define a throat 106 a at which a speed of thecombustion gases is sonic. The nozzle 106 defines a converging section106 b upstream of the throat 106 a and a diverging section 106 cdownstream of the throat 106 a. The nozzle 106 has an inlet 106 dfluidly connected to an outlet 104 a of the combustion chamber 104 andan outlet 106 e in fluid communication with an environment E outside thecombustion chamber 104. A cross-sectional area of the nozzle 106 takenon a plane normal to a longitudinal axis L of the rocket engine 102decreases from the inlet 106 d of the nozzle 106 to the throat 106 a andincreases from the throat 106 a to the outlet 106 e of the nozzle 106.

The rocket engine system 100 includes an oxidizer reservoir 112configured to contain the oxidizer. In a particular embodiment, theoxidizer is hydrogen peroxide (H₂O₂). The oxidizer may be a solutioncontaining 90% of hydrogen peroxide. Alternatively, the oxidizer may be,for instance, nitrous oxide, gaseous oxygen, liquid oxygen, nitrogentetroxide, nitric acid. The oxidizer is may be any conventional oxidantused for solid fuel, such as O₂, H₂O₂, HNO₃. Any suitable oxidizer maybe used without departing from the scope of the present disclosure. Thesolid propellant 110 may be, for instance, polyethylene, poly-methylmethacrylate, poly-vinyl chloride, hydroxyl terminated poly-butadiene,paraffin wax. More detail about the solid propellant 110 are presentedherein below.

Since the rocket engine system 100 is used in space, gravity may not besufficient for inducing a flow of the oxidizer in the combustion chamber104. In the embodiment shown, the rocket engine system 100 has a gasreservoir 114 containing a high-pressure gas, which may be helium. Thepressure of the gas contained in the gas reservoir 114 is such that itinduces a flow of the oxidizer contained in the oxidizer reservoir 112,from the oxidizer reservoir 112 to an inlet 104 b of the combustionchamber 104 of the rocket engine 102.

To allow the oxidizer to chemically react to yield oxygen and anothercomponent, the flow of oxidizer exiting the oxidizer reservoir 112passages through a catalyst 116. In the embodiment shown, the catalyst116 has an inlet 116 a for receiving the oxidizer and an outlet 116 bfor outputting the catalyzed oxidizer. The inlet 116 a of the catalyst116 is fluidly connected to the oxidizer reservoir 112. The outlet 116 bof the catalyst 116 is fluidly connected to the combustion chamber 104.The catalyst may contain a mesh 116 c coated with a suitable material,which may be silver. The mesh 116 c may be a mesh of SS304 stainlesssteel. The coating may be made of silver, rare metals, noble metals,ceramics, palladium, ruthenium, manganese oxide, oxides, iron salts, andso on. The material of the coating may maintain its structural andmechanical properties at temperatures of least at 660 degrees Celsius.The mesh 116 c may be, for instance, metallic wires with coating,ceramic pellets with catalyst sintered, 3D printed porous matrix withcoating.

In the embodiment shown, the oxygen used is obtained from a solution ofabout 70% of hydrogen peroxide, preferably about 90% and up to 99%, alsoreferred to as HTP. When contacting a material of the catalyst 116, theHTP is catalyzed and converted in gaseous oxygen and superheated watersteam (about 660 degrees Celsius). Using catalyzed HTP instead of liquidHTP may offer some advantages: improved combustion efficiency; improvedmixing with fuel; and presence of supersonic choke at injector, whichmay ensure improved combustion stability and decreased pressureoscillations in the combustion chamber 104. Moreover, since thecatalyzed HTP is hot, the rocket engine 102 may not need an independentheat source for ignition. In other words, the disclosed rocket engine102 may be free of an igniter. This may result in a simplified designcompared to a configuration having an igniter. Using the catalyzed HTPfor igniting the fuel may allow to relight the rocket engine 102 whenflying.

The gas reservoir 114, the oxidizer reservoir 112, the catalyst 116, andthe rocket engine 102 are fluidly connected with suitable conduits 118.In the embodiment shown, the pressure of the gas contained in the gasreservoir 114 induces a flow of the oxidizer from the oxidizer reservoir112, the oxidizer passages through the catalyst 116 where it reacts togenerate water and oxygen in gaseous phase. The catalyzed HTP isinjected in the combustion chamber 104 and ignites the solid propellantand generate combustion gases in the combustion chamber 104. Thecombustion gases are the accelerated through the nozzle 106 and expelledto the environment E to propel a vehicle equipped with the rocket engineassembly 100.

Still referring to FIG. 1, an injector manifold 120 is provided upstreamof the combustion chamber 104 and downstream of the catalyst 116relative to a flow of the oxidizer within the conduits 118. The injectormanifold 120 distributes the oxidizer in a suitable manner within thecombustion chamber 104. The injector manifold 120 is preferablywatertight as to limit leakage of the oxidizer and/or of the combustiongases. The injector manifold 120 is designed as to be able to withstandmechanical and thermal stresses. An injector plate 122 is locateddownstream of the injector manifold 120. The injector plate 122 may beprovided in the form of a circular plate defining a plurality ofapertures therethrough configured to allow passage of the oxidizerexiting the manifold 120. The injector manifold 120 has an inlet 120 afluidly connected to the outlet 116 b of the catalyst 116 and an outlet120 b fluidly connected to an inlet 122 a of the injector plate 122. Theinjector plate 122 has an outlet 122 b in fluid communication with thecombustion chamber 104. Different configurations of the injectormanifold 120 and the injector plate 122 are presented herein below.

The injector plate 122 may be annularly shaped and located at theconvergent section 106 b of the nozzle 106. The injector plate 122 maybe located anywhere along the combustion chamber 104 and nozzle 106. Insome cases, the injector plate 122 is a circular plate located upstreamof the combustion chamber 104 or an annular plate located proximate thenozzle 106. The injector plate 122 is preferably located in such a wayas to dispose the injector manifold 120 as close as possible from alocation where the injection of the catalyzed HTP is desired, such asclose to the propellant, to minimize pressure drops.

Many configurations of the solid propellant 110 within the rocket engine102 are possible. FIGS. 1a to 6c described below illustrate a pluralityof configurations of the solid propellant in the combustion chamber 104and show how the combustion of the solid propellant alternate theconfigurations of the solid propellant.

Referring to FIG. 1a , a rocket engine in accordance with one embodimentis generally shown at 10. The rocket engine 10 includes a housing 12that defines a combustion chamber 14 therein. The housing 12 has alongitudinal axis L. The housing 12 defines an inlet 12 a and an outlet12 b of the combustion chamber 14. A flow passage 16 extends from theinlet 12 a to the outlet 12 b. The inlet 12 a is fluidly connected to asource of an oxidizer, which may be an oxidizer tank (not shown) thatmay be part of the rocket engine 10. The outlet 12 b is fluidlyconnected to, or opens to, an environment E outside the combustionchamber 14 for expelling combustion gases generated within thecombustion chamber 14.

The housing 12 of the rocket engine 10 further defines a nozzle 12 caxially between the outlet 12 b and the combustion chamber 14. In theembodiment shown, the nozzle 12 c is a convergent-divergent nozzle andis used to accelerate the combustion gases generated within thecombustion chamber 14 to supersonic speeds. The nozzle 12 c includes athroat 12 d. The throat 12 d of the nozzle 12 c is where across-sectional area of the nozzle 12 c taken on a plane normal to thelongitudinal axis L is the smallest.

The housing 12 contains a first fuel 18 and a second fuel 20 bothcontained within the combustion chamber 14. The first fuel 18 has afirst solid propellant and the second fuel 20 has a second solidpropellant. A regression rate of the first solid propellant is greaterthan that of the second solid propellant. In a particular embodiment,the first fuel 18 is made of the first solid propellant and the secondfuel 20 is made of the second solid propellant. In a particularembodiment, a ratio of the regression rate of the first solid propellantover the regression rate of the second solid propellant ranges from 1.25to 30. In a particular embodiment, the rocket engine may include morethan two fuels differing from each other by their respective regressionrate.

Herein, regression rate means a rate at which a solid propellant isconsumed. The regression rate may be expressed in length units by timeunits (e.g., mm/s). The regression rate is also referred to as the burnrate and is the rate at which fuel can be induced to vaporize or ablateoff so it can participate in the combustion process and contribute torocket thrust.

Various possible arrangements of the first and second fuels arepossible. A few are described herein below with reference to FIGS. 1a ,2 a, 3 a, 4 a, 5 a, and 6 a. It is understood that the disclosedarrangements are only a few possibilities and the present disclosureshould not be limited by these arrangements. Other arrangements arecontemplated.

Still referring to FIG. 1a , the first fuel 18 and the second fuel 20are axially offset relative to the longitudinal axis L. As shown, thefirst and second fuels 18, 20 are located upstream of the nozzle 12 crelative to a direction of the combustion gases in the flow passage 16.In the embodiment shown, the first fuel 18 includes two annular disks 18a and the second fuel 20 includes three annular disks 20 a. An axialthickness of the annular disks 18 a of the first fuel 18 taken along thelongitudinal axis L is greater than that of the annular disks 20 a ofthe second fuel 20. As shown, the annular disks 18 a of the first fuel18 and the annular disks 20 a of the second fuel 20 are disposed inalternation along the longitudinal axis L. Stated otherwise, each of thetwo annular disks 18 a of the first fuel 18 is located between, orsandwiched, between two of the three annular disks 20 a of the secondfuel 20.

As shown, the flow passage 16 extends through apertures defined throughthe first and second fuels 18, 20. In the embodiment shown, peripheralsurfaces of the apertures defined through the first and second fuels 18,20 bounds the flow passage 16. This might allow the oxidizer to contactthe first and second fuels when it flows from the inlet 12 a to theoutlet 12 b within the flow passage 16.

Referring now to FIGS. 1a to 1c , the regression of the first and secondfuels 18, 20 as a function of time is illustrated. As the first solidpropellant of the first fuel 18 has a regression rate greater than thatof the second solid propellant of the second fuel 20, the first fuel 18is consumed more rapidly than the second fuel 20. In the embodimentshown, annular pockets 22 are formed following the more rapidconsumption of the first fuel 18. The annular pockets are bounded byannular surfaces defined by the annular disks 20 a of the second fuel22. Dimensions of these annular surfaces increase with time. In aparticular embodiment, the second fuel 20 acts like a diaphragm andincreases efficiency of the combustion, while contributing to the totalimpulse of the rocket engine 10, as it consumes itself.

Referring to FIGS. 2a to 2c , another embodiment of a rocket engine isgenerally shown at 200. For the sake of conciseness, only elements thatdiffer from the rocket engine 10 of FIG. 1a are described below.

In the embodiment shown, the first and second fuels 218, 220 areradially offset from each other relative to the longitudinal axis L ofthe housing 12. In the embodiment shown, the first and second fuels 218,220 are concentric tubes 218 a, 220 a; the second fuel 220 beingdisposed around the first fuel 218. In other words, the first fuel 218is located radially inwardly to the second fuel 220 relative to thelongitudinal axis L.

In the depicted embodiment, an axial length of the second fuel 220 isgreater than that of the first fuel 218 such that the second fuel 220protrudes beyond the first fuel 218 toward the outlet 12 b of thecombustion chamber 14. A radial thickness relative to the longitudinalaxis L of the first and second fuels 218, 220 may be equal beforecombustion starts.

In a particular embodiment, the second fuel 220 ensures film cooling.Film cooling occurs by insulating the surface to be protected (e.g., thehousing 12) from the rapidly flowing hot propellant gases by interposinga thin film of a cooling liquid fuel as it consumes itself along thesurface, flowing concurrently with the hot gases, to absorb and carryaway all or a portion of the total convective heat flux from the hotgases. In other words, the second fuel 220 is located adjacent thehousing 12 and radially between said housing 12 and the first fuel 218.As the second fuel 220 is consumed, it creates a flow of combustiongases in a vicinity of the housing 12. This flow might allow to cool thehousing 12.

Referring now to FIGS. 3a to 3c , another embodiment of a rocket engineis generally shown at 300. For the sake of conciseness, only elementsthat differ from the rocket engine 10 of FIG. 1a are described below.

As for the embodiment of FIG. 1, the first and second fuels 318, 320 areaxially offset relative to the axis L. The second fuel 320 is locatedaxially between the first fuel 318 and the outlet 12 b of the combustionchamber 14. In other words, the first fuel 318 is upstream of the secondfuel 320 relative to a direction of the combustion gases within the flowpassage 16.

In the embodiment shown, the first fuel 318 is an annular disk 318 a andthe second fuel 320 is an annular disk 320 a. In the depictedembodiment, dimensions of the annular disks 318 a, 320 a are the samebefore the combustion starts.

Referring to FIGS. 1a, 2a, and 3a , as the oxidizer is injected via theinlet 12 a of the combustion chamber 14, it contacts the first andsecond fuels 18, 20, 218, 220, 318, 320 that bound a portion of the flowpassage 16. As shown, the portion of the flow passage 16 correspond toperipheral surfaces of aperture defined by the first and second fuelsthrough which the flow passage 16 extends. As it is consumed, diametersof the peripheral surfaces increase thereby increasing a surface contactarea between the oxidizer and the first and second fuels. In otherwords, a dimension of the flow passage 16 increases with time as thefirst and second fuels are consumed. When the pockets 22 are created,they further increase the surface contact area between the oxidizer andthe first and second fuels.

Referring now to FIGS. 4a to 4c , another embodiment of a rocket engineis generally shown at 400. For the sake of conciseness, only elementsthat differ from the rocket engine 10 of FIG. 3a are described below.

In the embodiment shown, the first fuel 418 is a rod 418 a and thesecond fuel 420 is a tube 420 a disposed around the rod 418 a. Thesecond fuel 420 is disposed radially outwardly of the first fuel 418. Inother words, the first fuel 418 is surrounded by the second fuel 410. Ina particular embodiment, the second fuel 420, as it is consumed, ensuresfilm cooling.

The second fuel 420 protrudes axially beyond the first fuel 418 such asto define a pocket 422. The pocket 422 extends radially up to the secondfuel 420 and axially up the first fuel 418. As shown more clearly onFIGS. 4b and 4c , an axial depth of the pockets 422 relative to thelongitudinal axis L increases as the first fuel 418 is consumed.

In the embodiment shown, the inlet 412 a of the combustion chamber is atleast one aperture 412 a′ defined through the housing 412. The at leastone aperture 412 a′ may include a plurality of aperturescircumferentially distributed around the longitudinal axis L. Theaperture 412 a′ may be located axially between the first and secondfuels 418, 420 and the outlet 412 b of the combustion chamber 14. Theaperture 412 a′ may extend solely along a radial direction relative tothe longitudinal axis L. Alternatively, the aperture 412 a′ may extendalong both the radial direction and a circumferential direction relativeto the longitudinal axis L. The aperture 412 a′ may induce a swirl tothe oxidizer that is injected therethrough.

Referring now to FIGS. 5a to 5c , another embodiment of a rocket engineis generally shown at 500. For the sake of conciseness, only elementsthat differ from the rocket engine 10 of FIG. 3a are described below.

In the embodiment shown, the first fuel 518 is a first tube 518 a. Thesecond fuel 520 includes a rod 520 a and a second tube 520 b. The rod520 a is located within the first tube 518 a and both are located withinthe second tube 520 b. In other words, the first tube 518 a of the firstfuel 518 is located radially between the rod 520 a and the second tube520 b of the second fuel 520.

Thicknesses of the first and second tubes 518 a, 520 b taken in a radialdirection relative to the axis L may be equal or different.

In the embodiment shown, the inlet 512 a is at least one aperture 512 a′similar to the aperture 412 a′ described with reference to FIG. 4 a.

As shown more clearly on FIGS. 5b and 5c , annular pockets 522 arecreated by the combustion of the first fuel 518 at a rate greater thanthat of the second fuel 520. The creation of the annular pockets 522might allow to increase an area of the first fuel 518 that is in contactwith the oxidizer. This might enhance the production of the combustiongases and, hence, of the thrust generated by the engine.

In a particular embodiment, the average mass flow rates of the first andsecond fuels 18, 20 yields an optimal oxidizer to fuel (O/f) ratio, thusincreasing the overall yield of the combustion.

Referring now to FIGS. 6a to 6c , another embodiment of a rocket engineis generally shown at 600. For the sake of conciseness, only elementsthat differ from the rocket engine 10 of FIG. 3a are described below.

In the embodiment shown, the first fuel 618 includes a rod 618 a and afirst tube 618 b and the second fuel 620 includes a second tube 620 aand a third tube 620 b. The position of each of those components fromradially outward to radially inward relative to the longitudinal axis Lis as follows: the second tube 620 a of the second fuel 620, the firsttube 618 b of the first fuel 618, the third tube 620 b of the secondfuel 620, and the rod 618 a of the first fuel 618. Thicknesses of thesecond and third tubes 620 a, 620 b of the second fuel 620 may be lessthan that of the first tube 618 b of the first fuel 618.

As the first fuel 618 is consumed at a greater rate than the second fuel620, a central pocket 622 a and an annular pocket 622 b are created. Thecentral pocket 620 a is separated from the annular pocket 620 b by thesecond tube 618 b of the first fuel 618. As aforementioned, thosepockets might allow to increase an area of the first fuel 518 that is incontact with the oxidizer. This might enhance the production of thecombustion gases and, hence, of the thrust generated by the engine.

Referring to FIGS. 4a, 5a, and 6a , as the oxidizer is injected via theapertures 412 a′, it contacts the first and second fuels 418, 420, 518,520, 618, 620 that bound a portion of the flow passage 16. As shown, theportion of the flow passage 16 correspond to end surfaces of the firstand second fuels. Depending on the configuration, the end surfaces ofthe first and second fuels are circular or annular surfaces. As it isconsumed, axial lengths relative to the longitudinal axis L of the firstand second fuels decrease. A distance taken along the longitudinal axisL between the apertures 412 a′ and the first and second fuels increasesas they get consumed. As the first and second fuels do not get consumedat the same rate, a surface contact area between the first and secondfuels and the oxidizer increases. These additional surfaces that becomein contact with the oxidizer are created by the formation of thepockets. In other words, the additional surfaces are the surfaces thatbound the pockets 422, 522, 622 a, 622 b that are created by thecombustion of the first and second fuels.

In a particular embodiment, such a change of fuel geometry during burnallows a controlled increase of combustion surface, thus modulatingthrust curve during flight

In a particular embodiment, the first and second fuels 18, 20 have highvolumetric specific impulse that might allow a major volume reduction ofa tank containing the oxidizer and of the combustion chamber 14. Thismight increase structural margins of a vehicle equipped with thedisclosed engine 10 (which increases stage mass ratio). This increase ofstage mass ratio might compensate for the mass losses required forpressure-fed pressurant gas and the size of the pressurant tank.

Compared to solid engines, the disclosed rocket engines might surpassthe specific impulse of other existing propulsion systems discussedherein above. The disclosed fuels might be inherently safer to transportand operate. The fuel on its own might be stable and non-toxic and thecombustion might only occur when the oxidizer is injected into thecombustion chamber.

Overall, hybrid propulsion systems herein disclosed might differentiatethemselves from competing liquid engines by being far simpler andcheaper and having a higher volumetric specific impulse. They mightdifferentiate themselves from solid engines by being far safer toproduce and handle. And they might differentiate themselves fromstandard hybrids by having the requisite performance for orbital launchvehicles. The technology might also maintain a high degree offlexibility, meaning that it might be possible for it to be integratedinto several alternative launch solutions (air launch, balloon launch,etc.).

Having at least two kinds of solid propellant that differ by theregression rate might allow the ability to yield an optimal oxidizer tofuel ratio during the total duration of the burn thus achieving idealcharacteristic velocity. It might further offer the ability to alter theshape of the fuel during the combustion phase, obtaining some coolingcharacteristics.

For operating the rocket engines, the oxidizer is received within thecombustion chamber 14; the first and second fuels are exposed to thereceived oxidizer; and the combustion gasses created by a reactionbetween the received oxidizer and the fuels are expelled.

In the embodiment shown in FIGS. 1a, 2a, 3a , receiving the oxidizerincludes receiving the oxidizer via an inlet and expelling thecombustion gasses includes expelling the combustion gasses via anoutlet; the first and second fuels being located axially between theinlet and the outlet relative to a longitudinal axis of the rocketengine.

In the embodiment shown in FIG. 1a, 2a, 3a , receiving the oxidizerincludes receiving the oxidizer along an axial direction relative to alongitudinal axis of the rocket engine.

In the embodiments shown in FIGS. 4a, 5a, 6a , receiving the oxidizerincludes receiving the oxidizer in a direction having a radial componentrelative to the longitudinal axis L of the rocket engine. In aparticular embodiment, receiving the oxidizer in the direction havingthe radial component further comprises receiving the oxidiser in thedirection further having a circumferential component relative to thelongitudinal axis L.

Referring now to FIG. 7, another embodiment of a rocket engine is showngenerally at 700. In the embodiment shown, the fuel is a solidpropellant 718 in the form of a disk located within the combustionchamber 704 of the engine 700. Alternatively, the propellant may be atube defining a central passage 718 a.

As shown in FIG. 7, the injector manifold 720 and the injector plate 722are annular and extends circumferentially relative to the enginelongitudinal axis L all around the convergent section 706 b of thenozzle 706. The catalyst 716 is annular and located radially between theinjector plate 722 and the injector manifold 720.

In the embodiment shown, the catalyst catalyst 716 defines an inlet 716a and an outlet 716 b. Both of the inlet 716 a and the outlet 716 b ofthe catalyst catalyst 716 may be annular. In the embodiment shown, theinlet and outlet 716 a, 716 b of the catalyst correspond to externalsurfaces of the mesh 716 c. One or more of those external surfaces maybe in contact with the injector manifold 720 and another one of thosesurfaces may be in contact with the injector plate 722. The injectorplate 722 may define a portion of a wall 706 f of the nozzle 706.

In the depicted embodiment, the catalyst 716 is triangular when seen ina cross-section taken on a plane containing the longitudinal axis L ofthe engine 700. The catalyst 716 may include three annular surfaces; oneof the three surfaces may define the outlet 716 b of the catalyst 716;another one of the three surfaces may define a portion of the inlet 716a of the catalyst 716 and may be a substantially cylindrical surface;the remaining one of the three surfaces may define a remainder of theinlet 716 a of the catalyst 716 and may be have the shape of an annulardisk. In other words, the oxidizer may penetrate the catalyst 716 bothin an axial direction relative to the longitudinal axis L toward thesolid propellant 718 and in a radial direction toward the longitudinalaxis L. Other configurations are contemplated.

Referring to FIG. 8, in the embodiment shown, the injector plate 722defines a plurality of apertures 722 a that are circumferentiallydistributed about the longitudinal axis L. Several shapes for theapertures 722 a are considered such as, round, oblong, oval, and so on,or any suitable combinations thereof. Any suitable shape may be used.The apertures may have central axes 722 b oriented to inject thecatalyzed HTP with a circumferential component relative to thelongitudinal axis L. In other words, the apertures 722 a of the injectorplate 722 may extend from inlets 722 c to outlets 722 d; the outlets 722d being circumferentially offset form the inlets 722 c. Such aconfiguration may allow to inject the catalyzed HTP with a swirl.Although four apertures 722 a are shown, more or less apertures may beused without departing from the scope of the present disclosure. Theapertures 722 a, further to extend in a circumferential directionrelative to the longitudinal axis L, may extend in an axial and/or aradial direction relative to the longitudinal axis L. In a particularembodiment, the apertures are designed such that a flow of the catalystis chocked at the injector plate. This may improve stability of thecombustion and may improve an efficiency of the catalyst.

Referring to FIG. 9, another embodiment of an injector plate, which maybe used with the rocket engine 700 described above with reference toFIG. 7, is shown generally at 822. The injector plate 822 definesapertures 822 a that are circumferentially distributed all around thelongitudinal axis L of the rocket engine 700. Several shapes for theapertures 822 a are considered such as, round, oblong, oval, and so on,or any suitable combinations thereof. Any suitable shape may be used. Inthe embodiment shown, the apertures 822 a have central axes 822 boriented axially relative to the longitudinal axis L. In other words,the inlets 822 c of the apertures 822 may be radially aligned with theoutlets 822 d of the apertures relative to the longitudinal axis L ofthe engine. The inlets and outlets 822 c, 822 d of the apertures 822 amay be circumferentially aligned relative to the longitudinal axis L ofthe rocket engine. The central axes 822 b of the apertures 822 a may beparallel to the engine longitudinal axis L. The apertures 822 a mayextend axially and circumferentially relative to the engine longitudinalaxis L and may be free of a radial component. Although two apertures 822a are shown, more or less apertures may be used without departing fromthe scope of the present disclosure. The apertures 822 a may bechamfered or have rounded edges. This may allow for better flowcirculation.

It is understood that any combinations of the injector plates 722, 822described above with reference to FIGS. 7a and 8 are contemplated. Forinstance, the apertures may extend axially, radially, and/orcircumferentially relative to the longitudinal axis L of the engine. Asaforementioned, the apertures may have different shapes and may beangled. Any suitable combinations of two or more shapes for theapertures may be used.

Referring now to FIG. 10, a rocket engine in accordance with anotherembodiment is generally shown at 900. The rocket engine 900 includes asolid propellant 918 in the form of a tube located within the combustionchamber 904 of the engine 900. The engine 900 includes a manifold 920,an injector plate 922, and a catalyst 916 between the manifold 920 andthe injector plate 922. The oxidizer is injected in the manifold 920,which is configured to distribute the oxidizer all around the catalyst916. The oxidizer then penetrates the catalyst 916 and is injected inthe combustion chamber 914 via the injector plate 922. The outlet of themanifold 920 is fluidly connected to the inlet 916 a of the catalyst916. The outlet 916 b of the catalyst 916 is fluidly connected to thecombustion chamber 904 via the injector plate 922.

In the embodiment shown, the catalyst 916 contains a mesh in the form ofa disk and located at an axial end of the combustion chamber 904. Thecatalyst 916 has an inlet 916 a and an outlet 916 b. The inlet and theoutlet 916 a, 916 b of the catalyst 916 correspond to external faces ofthe disk. For instance, the inlet 916 a may correspond to a firstcircular external face of the disk and to a cylindrical face of the diskwhereas the outlet 916 b may correspond to a second circular externalface of the disk opposed the first circular external face. In otherwords, and in the embodiment shown, the oxidizer may penetrate thecatalyst 916 axially and radially relative to the longitudinal axis L ofthe engine 900.

The oxidizer is injected in the manifold 920 via an inlet thereof anddistributed all around the inlet 916 a of the catalyst 916. The oxidizerthen passes through the catalyst 916 and exits the catalyst 916 via theoutlet 916 b thereof, and passes through the apertures of the injectionplate 922 before being injected in the combustion chamber 904. In theembodiment shown, the oxidizer is injected substantially axiallyrelative to the longitudinal axis L and within a hollow section definedby the propellant 918.

Referring to FIG. 11, the injector plate 922 is shown on a cross-sectiontaken on a plane containing the longitudinal axis L of the engine 900.As illustrated, the injector plate 922 includes a plurality of apertures922 a. Several shapes for the apertures 922 a are considered such as,round, oblong, oval, and so on, or any suitable combinations thereof.Any suitable shape may be used. The apertures 922 a fluidly connect thecombustion chamber 904 to the manifold 920 and the catalyst 916. In theembodiment shown, the apertures 922 a have inlets 922 b and outlets 922c that are circumferentially aligned with one another. In other words,and in the depicted embodiment, the apertures 922 a extend axiallythrough the injector plate 922 relative to the longitudinal axis L ofthe engine 900. Stated differently, the apertures may extend parallel tothe longitudinal axis L of the engine 900. In such a case, the oxidizeris injected in the combustion chamber 904 in a substantially axialdirection relative to the longitudinal axis L.

Referring now to FIG. 12, another embodiment of an injector plate isshown generally at 1022. The injector plate 1022 has a substantiallydisk shape and defines a plurality of apertures 1022 a therethrough. Theapertures 1022 a have inlets 1022 b and outlets 1022 c. In theembodiment shown, the inlets 1022 b are circumferentially offset fromthe outlets 1022 c relative to the longitudinal axis L of the engine. Insuch a case, the oxidizer is injected in the combustion chamber in adirection being axial and circumferential relative to the longitudinalaxis L. Therefore, the injector plate 1022 may induce a swirl in theoxidizer. Although four apertures are shown, more or less apertures maybe used without departing from the scope of the present disclosure. Theswirl may allow for a better stability of the flow, a better surfaceflux, a better regression rate, longer residency time in the combustionchamber, and better combustion efficiency.

Referring now to FIG. 13, another embodiment of a rocket engine is showngenerally at 1100. For the sake of conciseness, only elements thatdiffer from the engine 900 described above with reference to FIGS. 10and 10 a are described herein below.

In the embodiment shown, the engine 1100 includes a catalyst 1116disposed between a manifold 1120 and an injector plate 1122. Thecatalyst 1116 has an annular shape and extends circumferentially allaround the longitudinal axis L of the engine 1110 and around acombustion chamber 1104 thereof. In the embodiment shown, the solidpropellant 1118 is located axially between the injector plate 1112 andthe nozzle 1106 relative to the longitudinal axis L.

The catalyst 1116 may have a square or rectangular shape when seen in across-section taken on a plane containing the longitudinal axis L of theengine 1100. The catalyst 1116 contains a mesh having three externalfaces defining an inlet of the catalyst, namely two annular faces andone cylindrical face in the embodiment shown, and one face, namely acylindrical face, defining an outlet of the catalyst.

The manifold 1120 may have a U-shape when seen in cross-section on aplane containing the longitudinal axis L of the engine 1100. Themanifold 1120, in the depicted embodiment, is annular and extends allaround the longitudinal axis L of the engine 1100. The manifold 1120 isconfigured to distribute the oxidizer on all of the three faces of themesh of the catalyst 1116 that define its inlet. In the present case,the manifold 1120 defines three injection faces 1120 a each beinglocated adjacent a respective one of the three faces of the meshdefining the inlet of the catalyst 1116.

Referring to FIG. 14, the injector plate 1122 is illustrated. Theinjector plate 1122 includes a plurality of apertures 1112 acircumferentially distributed around the longitudinal axis L of theengine 1100. The apertures 1122 a have inlets 1122 b and outlets 1122 c.The outlets 1112 c of the apertures 1122 a of the injector plate 1122are fluidly connected to the combustion chamber 1104 (FIG. 12). In theembodiment shown, the inlets 1122 b and the outlets 1122 c arecircumferentially offset from one another relative to the longitudinalaxis L of the engine. In such a case, the oxidizer is injected in thecombustion chamber 1104 in a direction being radial and circumferentialrelative to the longitudinal axis L. Therefore, the injector plate 1122may induce a swirl in the oxidizer. It is understood that the inlets andoutlets 1122 b, 1122 c of the apertures 1122 a may be axially offsetfrom one another relative to the longitudinal axis L such that theoxidizer, further to be injected with a circumferential component isinjected with an axial component, which may be oriented toward thepropellant 1118. Alternatively, the inlets and outlets 1122 b, 1122 c ofthe apertures 1122 a may be axially aligned relative to the longitudinalaxis L.

Referring now to FIG. 15, a section of a rocket engine is showngenerally at 1200. The section 1200 may define an upstream end, such asa header, of the rocket engine and a portion of a combustion chamber1204 thereof. In the embodiment shown, the rocket engine includes anintegrated manifold-catalyst-injector plate, referred to herein belowhas an Integrated Injector Assembly (IIA) 1230.

The IIA 1230 has an external wall 1230 a, an internal wall 1230 blocated radially inwardly of the external wall 1230 a relative to thelongitudinal axis L, and an annular cavity 1230 c shown in dashed lineand located radially between the external and internal walls 1230 a,1230 b. Both of the external and internal walls 1230 a, 1230 b areannular and extend circumferentially around the longitudinal axis L. TheIIA 1230 has an inlet defined by a plurality of circumferentiallydistributed apertures 1230 d, also referred to as oxidizer inlets orliquid peroxide inlets, extending through the external wall 1230 a andan outlet defined by a plurality of circumferentially distributedapertures 1230 e, also referred to as catalyzed injection ports, orcatalyzed peroxide injection ports, extending through the internal wall1230 b of the IIA 1230.

In the embodiment shown, the catalyst 1216 is located within the annularcavity 1230 c radially between the external and internal walls 1230 a,1230 b. In the embodiment shown, the annular cavity 1230 c is dualpurpose has it may allow the oxidizer to be circumferentiallydistributed all around the longitudinal axis L while being catalyzed bythe catalyst 1216 since the catalyst 1216 is located inside the annularcavity 1230 c. In other words, the IIA 1230 performs the function of amanifold, a catalyst, and of an injection plate. In the present case,the injection plate may be considered as the internal wall 1230 b.

In the embodiment shown, the apertures 1230 e defined through theinternal wall 1230 b defines exit flow axes C, each having acircumferential component relative to the longitudinal axis L to inducea swirl in the oxidizer as it is injected in the combustion chamber1204. In other words, the internal wall 1230 b and the apertures 1230 edefined therethrough act as a swirl injector. It is understood thatother configurations are contemplated. For instance, the apertures 1230e may be oriented radially and/or circumferentially or any suitableorientation. For instance, the internal wall 1230 b and apertures 1230 emay be coaxial, vortex, doublets, pintle injectors and so on.

There is disclosed herein a catalyst embedded within the injectionmanifold, all in one part (e.g., the IIA 1230). The IIA 1230 may bemanufactured using additive manufacturing. This component, whichincludes the injector as well, can be integrated anywhere along thecombustion chamber. It is understood that other configurations arepossible. For instance, the IIA may have a disk shape instead of anannular shape; the catalyst may be located axially between an externalwall and an internal wall, which may be provided in the form of disks.Suitable apertures may be defined through the external and internalwalls for distributed the oxidizer in the catalyst and for distributingthe catalyzed oxidizer in the combustion chamber 1204.

Referring now to FIG. 16, a rocket engine in accordance with anotherembodiment is shown generally at 1300. For the sake of conciseness, onlyelements that differ from the rocket engine 102 described herein abovewith reference to FIG. 1 are described herein below. The rocket engine1300 has a combustion chamber 1304 and a nozzle 1306 in fluidcommunication with said chamber 1304. In the embodiment shown, theinjector plate 1322 is located at the converging section 1306 a of thenozzle 1306, but other configurations are contemplated. Anyconfigurations described above with reference to FIGS. 7-15 may be used.

Typically, a rocket engine encompassing a single solid propellant, beinguniform and homogeneous, will burn very fast initially, then would startto burn slower and slower as the flame front moves along a regressiondirection. This may be explained by an increase of a volume of thecombustion chamber with time as the solid propellant is graduallydepleted. Therefore, with time, an increased volume of the combustionchamber may imply a lower concentration of the oxidizer in thecombustion chamber, which may yield a lower regression rate. It may bepossible to vary a mass flow rate of the oxidizer (e.g. to increase themass flow rate) as the propellant is depleted to limit a decrease in theconcentration of the oxidizer.

Herein, a regression direction is a direction along which the solidpropellant is being depleted. In other words, the regression directionrepresents a direction along which a dimension of the solid propellantdecreases during burn. For instance, if the solid propellant is in theform of a tube defining a central passage, the regression direction maybe a radial direction as a radial thickness of the tube decreases as thepropellant is burned away during the combustion process. In such a case,a dimension of the central passage may increase with time during thecombustion process. As another example, if the solid propellant is inthe form of a solid cylinder having an axial circular end face exposedto the oxidizer, the regression direction may be an axial direction as alength of the solid propellant decreases during the combustion process.Other configurations are contemplated.

In the embodiment shown, the combustion chamber 1304 contains aplurality of fuels each having a solid propellant. In the embodimentshown, the combustion chamber 1304 contains three fuels, namely a first,a second, and a third fuel 1318 a, 1318 b, 1318 b that are axiallystacked one next to the other relative to the longitudinal axis L of theengine 1300. These three fuels 1318 a, 1318 b, 1318 c are, in thepresent embodiment, cylinders or disks. Each of these three fuels 1318a, 1318 b, 1318 c is made of a solid propellant differing from that ofthe others by their regression rates. This may be achieved, forinstance, by having the three fuels differing by one or more rheologicalproperties.

The regression direction is illustrated on FIG. 16 by arrow R. The firstfuel 1318 a has an axial end face 1318 a ₁ exposed to the oxidizer beinginjected in the combustion chamber 1304 by the injector plate 1322.Therefore, during the combustion process, a length taken along thelongitudinal axis L of the engine, decreases with time. At some point,the first fuel 1318 a is totally consumed and the second fuel 1318 bstarts to be burned again along the regression direction R. When thelength of the second fuel 1318 b is zero, meaning that the second fuel1318 b is consumed in entirety, the third fuel 1318 c starts to beconsumed and its length decreases along the regression direction R untilno more fuel remains within the rocket engine.

In the present case, as the fuel stacking burns, the impact of thevariation of the oxidizer flux, which may cause the fuel to burn slowerover time, may be compensated by the variation of the regression ratesof the three fuels 1318 a, 1318 b, 1318 c. In the embodiment shown, therheological property that is different between the three fuels 1318 a,1318 b, 1318 c is the viscosity. Varying the viscosity may guarantee asubstantially constant fuel mass flow. This constant fuel mass flow maycontribute in maintaining a substantially stable oxidizer-to-fuel ratio,which in turn may allow the disclosed rocket engine 1300 to have betterperformances than a rocket engine having a single solid propellant ofuniform viscosity.

It is understood that the three fuels 1318 a, 1318 b, 1318 c may differby any other rheological properties, alternatively or in combination.Those other properties may be, for instance, the crystallinity of thefuel. Other rheological properties may be, for instance, the density,mechanical properties vs. temperature, fusion temperature, glasstransition temperature. It is understood that those rheologicalproperties are considered before the solid propellant enters a reactingphase. In a particular embodiment, an increase of the viscosity impliesa decrease in the regression rate. The regression rate may be verysensitive to a variation in the viscosity.

Referring now to FIG. 17, a rocket engine in accordance with anotherembodiment is shown generally at 1400. For the sake of conciseness, onlyelements that differ from the rocket engine 102 described herein abovewith reference to FIG. 1 are described herein below. The rocket engine1400 has a combustion chamber 1404 and a nozzle 1406 in fluidcommunication with said chamber 1404. In the embodiment shown, theinjector plate 1422 is located at the converging section 1406 a of thenozzle 1406, but other configurations are contemplated. Anyconfigurations described above with reference to FIGS. 7-15 may be used.

In the embodiment shown, the rocket engine 1400 includes a single solidpropellant 1418 that may burn along the regression direction R, whichmay correspond to an axial direction relative to the longitudinal axis Lof the engine. The solid propellant 1418 may define a gradient of arheological property; the gradient being in the regression direction R.In other words, the solid propellant 1418 may have a rheologicalproperty that varies through the propellant along the regressiondirection R. In the depicted embodiment, the solid propellant 1418 has aviscosity that decreases from a first end 1418 a located adjacent thenozzles 1406 to a second end 1418 b opposed to the first end 1418 a. Thefirst end 1418 a is the one exposed to the oxidizer. One or morerheological property(ies) may vary throughout the solid propellant alongthe regression direction R. The one or more rheological property(ies)may vary monotonically along the regression direction R, linearly,exponentially, or any suitable type of variation may be used.

The solid propellant 1418 may be manufactured using additivemanufacturing, and/or modern polymer casting techniques. The solidpropellant 1418 may have a rheological property (e.g., the viscosity)that vary in the axial direction relative to the longitudinal axis Land/or in the radial direction. Manufacturing techniques, such asadditive manufacturing, and/or modern polymer casting techniques, mayallow to cast a single solid fuel, but with a variable viscositythroughout its length (or radius, depending on the fuel's configurationwithin the combustion chamber). This can be achieved using differentmethods, namely by varying the thermal curing cycle of the cast over itslength (or radius). Another way to yield such a variation of viscositywithin the length of the fuel 1418 may be to pre-mold the fuel with arheological additive incorporated within, as to add less and less of thesaid additive the more a distance from the nozzle 1406 increases.Rheological additives may be, for instance, metal powders, oxides ofmetal powders, other miscible substances such as polymers, fibers, orcomposite types of powders, filler/inert additives. The compositepowders may be, for instance, carbon nanotubes, glass fibers, polyamide,etc.

The disclosed rocket engine 1400 having the solid propellant 1418characterized by a gradient in one or more rheological properties mayallow to obtain a substantially constant oxidizer-to-fuel ratio and mayallow to obtain a substantially constant mass flow.

Referring now to FIG. 18, a convergent-divergent nozzle in accordancewith another embodiment is shown generally at 1500. The nozzle 1500 maybe used in conjunction with any of the rocket engines disclosed aboveand described with reference to FIGS. 1-17. The nozzle 1500, asdiscussed above, has an inlet 1502, an outlet 1504, a converging section1506 and a diverging section 1508.

Typically, means are provided for controlling a direction of a thrustgenerated by the rocket engine for guiding the rocket. In some cases, atechnique known as ‘gimballing’ is used. In such a technique, one ormore actuators, such as hydraulic actuators, are used to pivot thenozzle to change a direction via which the combustion gases are expelledand therefore to change a direction of flight of the rocket. Thosehydraulic actuators use hydraulic pumps and lines to supply theactuators with hydraulic fluid. For light-to-medium class launchvehicles, this technique may be overly complex, heavy, and expensive.

Herein, it is proposed to rely on a bypass flow, bled from the mainengine, in the form of a gas or liquid, to effect a deflection in athrust vector V. The concept, broadly known as fluidic thrust vectoring(FTV), may allow developing low-cost space vehicles because it may besubstantially free of moving parts; it may be lightweight, and it mayoffer fast response times. Three distinct classes of FTV exist:shock-vector, counter flow, and throat shift. In the embodiment shown,the nozzle 1500 uses shock-vector control (SVC).

In SVC, a secondary fluid is injected in the diverging section 1508 ofthe nozzle 1500 where the flow is supersonic. The injected secondaryfluid may act as an obstacle that may be “foreseen” by the flow. Thismay result in a complex shock wave pattern that may create an asymmetricdistribution in wall pressure. The wall pressure, when integrated overthe nozzle area, may yield a deflection in the thrust vector V′ by anangle A1. Two distinct types of injection are possible in nozzle flows:(a) circular injection through an orifice and (b) cylindrical injectionthrough a slot. In some cases, the angle A1 may reach 4 degrees. Thebypass flow may be used to cause a deflection in the thrust vector,which may avoid using traditionally heavy hydraulics.

Still referring to FIG. 18, the nozzle 1500 includes a thrust vectorcontrol (TVC) device 1510 that may be operatively connected to thedivergent section 1508 of the nozzle 1500. The TVC device 1510 may beoperable to inject a fluid within a flow passage F of the nozzle 1500.The TVC device 1510 may inject the fluid in an asymmetric manner such asto alter an axisymmetry of the flow pattern within the nozzle 1500 tochange a direction of the force vector V.

The fluid that is injected may be contained within a reservoir 1512. Thereservoir 1512 may be the oxidizer reservoir for the rocket. In otherwords, the fluid injected in the divergent section 1508 of the nozzle1500 may correspond to the oxidizer injected in the combustion chamberof a rocket engine equipped with the disclosed nozzle 1500. The sameoxidizer reservoir may feed both of the combustion chamber and the TVCdevice 1510. Alternatively, the fluid may be contained in a dedicatedreservoir. The fluid may be the same as the oxidizer, may be a differentfluid than the oxidizer, may be in gaseous or liquid phase. The fluidmay be, for instance, HTP, or any other suitable fluid, such as anoxidizer or a fuel.

In the embodiment shown, the fluid is injected in the flow passage ofthe nozzle 1500 at the divergent section 1508 via at least one aperture1514. A plurality of apertures 1514 may be used and may becircumferentially distributed around a central axis C of the nozzle1500. In some cases, only one aperture 1514 may be used, for instance,when a rocket is equipped with more than one rocket engines each havingits respective nozzle. Typically, a minimum of three apertures 1514 areused. In some cases, more than 30 apertures are used. Typically, havingmore apertures may allow for a better granularity in the control of thedirection of the thrust vector V. The apertures 1514 may beequidistantly separated from one another. The aperture 1514 may belocated at from about 25% to about 70% of a length of the divergentsection 1508 along the central axis C from a throat 1516 of the nozzle1500. The apertures 1514 may be circular or may be slots extending in acircumferential direction relative to the central axis C. The positionof the apertures 1514, their number, their dimensions and so on may beoptimized as to allow a maximum deviation of the thrust vector V using aminimum volume of the fluid.

In the embodiment shown, the fluid reservoir 1512 is fluidly connectedto the apertures 1514 via valves 1518. The valves 1518 may beindependently operable from one another to allow injection of the fluidthrough the apertures 1514 independently from one another forcontrolling the angle of deflection A1 of the thrust vector V′. Suitableconduits 1511 are used to connect the oxidizer reservoir 1512 to thevalves 1518.

Referring to FIG. 15-16, a control system for the TVC 1510 is showngenerally at 1600. The control system 1600 includes a processing unit1602 operatively connected to a computer readable medium 1604. Thecontrol system 1600 may be operatively connected to one or more sensors1606, for instance a gyroscope. The control system 1600 may beoperatively connected to the valves 1518 for controlling their openingand closing. The computer readable medium 1604 may have instructionsstored thereon to obtain data from the sensor(s) 1606 and to determine arequired orientation of the thrust vector V′ to orient the rocket in adesired direction. The instructions may, based on the received data fromthe sensor 1606, calculate how to inject the fluid in the apertures1514. The control system 1600 may, for instance, determine which of thevalves 1518 to open, what mass flow rate of the fluid to inject, a timeduration of an injection of the fluid, and so on. The control system1600 may be able to determine when the rocket is in the desireddirection and close the valve to maintain the rocket in said direction.These calculations by the control system 1600 may be made continuouslyin real time to maintain the rocket in a desired orientation.

Referring back to FIG. 18, in operation, the temperatures of thecombustion gases circulating within the flow path F of the nozzle 1500may be very high and may affect proper operation of the valves 1518. Inthe depicted embodiment, the valves 1518 are fluidly connected to theapertures 1514 via respective conduits 1520. The conduits 1520 extend atleast partially radially away from a wall 1522 of the nozzle 1500 suchthat the valves 1518 are further away from combustion gases than if theconduits 1520 were not present.

In some cases, the temperature of the combustion gases in the flow pathF may be such that it is a challenge to maintain a watertight connectionbetween the conduits 1520 and the apertures 1514. In other words,mechanical connections, such as threads, or other types of connectionssuch as brazing and welding, may fail because of the high temperatures.

In the embodiment shown, at least the divergent section 1508 of thenozzle 1500 has a monolithic body 1524, which may be made of Inconel718™, Hastalloy™, or any suitable material able to withstandtemperatures of the combustion gases. The monolithic body 1524 maydefine an entirety of the nozzle 1500. A coating 1530, which may be madeof a composite material such as an ablative resin, and/or a phenolicresin, may be secured to the inner face 1522 a of the nozzle 1500. Thecoating 1530 may include fibers, such as carbon fibers and/or glassfibers. Although the coating 1530 is illustrated as being located solelyin the divergent section of the nozzle, it may cover the inner surfaceof an entirety of the nozzle. The monolithic body 1524 defines the wall1522 that extends circumferentially around the central axis C, definesthe apertures 1514 extending through a thickness T of the wall 1522, andfurther defines the conduits 1520. The conduits 1502 may protrudemonolithically away from the wall 1522. The body 1524, including thewall 1522 and the conduits 1520, may be made of a single block. In otherwords, the body 1524 may be made of a monolithic piece of material; saidmonolithic piece defining the wall 1522 and the conduits 1520.

In the embodiment shown, an entirety of the nozzle 1500 is made of themonolithic body 1524. However, it is understood that solely thedivergent section 1508 and the conduits 1520 may be made of a monolithicpiece of material and be secured to another body corresponding to theconvergent section 1506. The throat of the nozzle 1500 may bemanufactured with the convergent section 1506 or with the divergentsection 1508.

Having the conduits 1520 being monolithic with the wall 1522 may providefor a watertight connection at a junction between the conduits 1520 andthe wall 1522. Moreover, having the conduits 1520 being monolithic withthe wall 1522 may allow to increase a number of the conduits 1520without substantially increasing a risk of leakage.

Still referring to FIG. 18, a peripheral surface 1514 a of the apertures1514, or an inner surface of the conduit 1520, may be manufactured todefine a convergent divergent shape 1514 b having a throat 1514 c. Thismay allow to inject the secondary fluid at a greater speed than if acylindrical aperture were used. Manufacturing the body 1524 as amonolithic piece may allow to precisely control a shape of the apertures1514 to create the convergent divergent shape 1514 a.

In the embodiment shown, the nozzle 1500 includes a cooling system 1526that may be used to maintain a temperature of the wall 1522 of thenozzle 1500 within acceptable limits. In the embodiment shown, thecooling system 1526 includes a coolant passage 1528 that extendscircumferentially around the central axis C. The coolant passage 1528 islocated between inner and outer faces 1522 a, 1522 b of the wall 1522 ofthe monolithic body 1524. The coolant passage 1528 may be helicoidallyextending around the central axis C. Alternatively, a plurality ofcoolant passages may extend substantially axially along the central axisC and between the inner and outer surfaces 1522 a, 1522 b of thedivergent section 1508 of the nozzle 1500; manifolds may be fluidlyconnected to opposed ends of the plurality of conduits.

In the embodiment shown, the coolant passage 1528 is fluidly connectedat an inlet 1528 a thereof to the oxidizer reservoir 1512 and at anoutlet 1528 b thereof to the combustion chamber of the rocket engine. Asuitable conduit 1515 is used to connect the inlet 1528 a to theoxidizer reservoir 1512. The coolant passage 1528 may circulate theoxidizer from the oxidizer tank before the oxidizer is fed to thecombustion chamber. A temperature of the oxidizer may increase via itspassage within the coolant passage 1528 thereby cooling the wall 1522 ofthe nozzle 1500. The heated oxidizer may then be injected in thecombustion chamber. Alternatively, a dedicated coolant may circulatewithin the coolant passage 1528. Heating the oxidizer as such may allowto increase an entropy and enthalpy of the oxidizer and may allow toincrease a combustion efficiency compared to a configuration in whichthe oxidizer is not heated. Moreover, heating the oxidizer may allow toavoid using a catalyst since the oxidizer may be catalyzed via itspassage in the coolant passage 1528 without the need of a catalyst. Thenozzle 1500 may alternatively be manufactured without the coolingsystem.

The coating 1530 may be made of a material that is configured to beslowly ablated away with time during the combustion process of the solidpropellant within the combustion chamber of the rocket engine. Aregression rate of the material of the coating 1530 may be substantiallyless than that of the solid propellant(s) contained within thecombustion chamber of the rocket engine. The material of the coating1530, when exposed to hot combustion gases, may burn and generate a gasthat flows in a vicinity of, and parallel to, the inner surface 1522 aof the wall 1522 and may allow to film cool the wall 1522 of the nozzle1500 as the gas is pushed outside of the nozzle 1500 by the combustiongases circulating therethrough.

For manufacturing the divergent section 1508 of the nozzle 1500, themonolithic body 1524 may be manufactured using an additive manufacturingprocess to define the wall 1522 circumferentially extending about thecentral axis C and to define the conduits 1520 protruding away from thewall 1522.

The apertures 1514 may be manufactured during the additive manufacturingprocess. Alternatively, the apertures 1514 may be bored through the wall1522. Each of the apertures 1514 is concentrically aligned with arespective one of the conduits 1520. The coolant conduit or passage 1528may be manufactured within the thickness T of the wall 1522 via theadditive manufacturing process.

Embodiment disclosed herein includes:

A. A rocket engine comprising: a combustion chamber having a chamberinlet for receiving an oxidizer and a chamber outlet for expellingcombustion gases in an environment outside the combustion chamber; amanifold having a manifold inlet fluidly connectable to a source of theoxidizer and a manifold outlet; a catalyst having a catalyst inletfluidly connected to the manifold outlet and a catalyst outlet; and aninjector plate having a injector inlet fluidly connected to the catalystoutlet and an injector outlet fluidly connected to the chamber inlet.

B. An integrated injector assembly for injecting an oxidizer within acombustion chamber of a rocket engine, comprising an inner wall definingan inlet fluidly connectable to a source of an oxidizer, an outer wallspaced apart from the inner wall to define a cavity therebetween, theouter wall defining an outlet for injecting a catalyzed oxidizer withinthe combustion chamber, and a catalyst located within the cavity.

C. A method of supplying an oxidizer within a combustion chamber of arocket engine, comprising: receiving an oxidizer; distributing thereceived oxidizer while catalyzing the received oxidizer; and injectingthe catalyzed oxidizer within the combustion chamber.

Embodiments A, B, and C may include any of the following elements, inany combinations:

Element 1: the catalyst is located between the manifold and the injectorplate. Element 2: the injector plate includes a plurality of aperturesextending therethrough. Element 3: the plurality of apertures areoriented parallel to a longitudinal axis of the rocket engine. Element4: at least some of the plurality of apertures have aperture inletsbeing circumferentially offset from aperture outlets relative to alongitudinal axis of the rocket engine for creating a swirl in theoxidizer. Element 5: the catalyst has a mesh, the mesh coated withsilver. Element 6: the rocket engine includes a convergent-divergentnozzle located downstream of the combustion chamber, the injector plate,the manifold, and the catalyst being annular and extendingcircumferentially around a convergent section of theconvergent-divergent nozzle. Element 7: the injector plate includes aplurality of apertures extending therethrough. Element 8: the pluralityof apertures are oriented parallel to a longitudinal axis of the rocketengine. Element 9: at least some of the plurality of apertures haveaperture inlets being circumferentially offset from aperture outletsrelative to a longitudinal axis of the rocket engine for creating aswirl in the oxidizer. Element 10: the injector plate is a disk andwherein the catalyst includes a mesh being disk-shaped, the mesh havingat least two faces, one of the at least two faces located proximate theinjector plate and defining the outlet of the catalyst, the manifoldextending over the other of the at least to faces, the other of the atleast two faces defining the inlet of the catalyst. Element 11: theinjector plate extends circumferentially all around a longitudinal axisof the rocket engine, the catalyst being annular and extending aroundthe injector plate and located radially outwardly of the injector platerelative to a longitudinal axis of the rocket engine, the catalysthaving a mesh having at least two faces, one of the at least two faceslocated proximate the injector plate and defining the outlet of thecatalyst, the manifold extending over the other of the at least tofaces, the other of the at least two faces defining the inlet of thecatalyst. Element 12: the injector plate, the manifold, and the catalystare located at a convergent section of a convergent-divergent nozzle ofthe rocket engine. Element 13: the inner wall and the outer wall areannular and circumferentially extend all around a longitudinal axis, theinner wall located radially inwardly of the outer wall, the cavity andthe catalyst being annular and extending around the longitudinal axis.Element 14: the inlet includes a plurality of apertures extendingthrough the outer wall and circumferentially distributed around thelongitudinal axis. Element 15: the outlet includes a plurality ofapertures extending through the inner wall and circumferentiallydistributed around the longitudinal axis. Element 16: the plurality ofapertures have exit flow axes having each a circumferential componentrelative to the longitudinal axis. Element 17: distributing the receivedoxidizer includes distributing the received oxidizer within an annularcavity extending all around a longitudinal axis of the rocket engine,and catalyzing the received oxidizer includes circulating the receivedoxidizer through a mesh extending around the longitudinal axis.

In the present specification including claims, the term “about” meansthat a value may range from the value minus 10% of the value to thevalue plus 10% of the value. For instance, a value of about 10 impliesthat the value ranges from 9 to 11.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.For example, other shapes and arrangements of the fuels within thecombustion chamber are contemplated. More than two fuels may be used.Still other modifications which fall within the scope of the presentinvention will be apparent to those skilled in the art, in light of areview of this disclosure, and such modifications are intended to fallwithin the appended claims.

1. A rocket engine comprising: a combustion chamber having a chamberinlet for receiving an oxidizer and a chamber outlet for expellingcombustion gases in an environment outside the combustion chamber; amanifold having a manifold inlet fluidly connectable to a source of theoxidizer and a manifold outlet; a catalyst having a catalyst inletfluidly connected to the manifold outlet and a catalyst outlet; and aninjector plate having a injector inlet fluidly connected to the catalystoutlet and an injector outlet fluidly connected to the chamber inlet. 2.The rocket engine of claim 1, wherein the catalyst is located betweenthe manifold and the injector plate.
 3. The rocket engine of claim 1,wherein the injector plate includes a plurality of apertures extendingtherethrough.
 4. The rocket engine of claim 3, wherein the plurality ofapertures are oriented parallel to a longitudinal axis of the rocketengine.
 5. The rocket engine of claim 3, wherein at least some of theplurality of apertures have aperture inlets being circumferentiallyoffset from aperture outlets relative to a longitudinal axis of therocket engine for creating a swirl in the oxidizer.
 6. The rocket engineof claim 1, wherein the catalyst has a mesh, the mesh coated withsilver.
 7. The rocket engine of claim 1, wherein the rocket engineincludes a convergent-divergent nozzle located downstream of thecombustion chamber, the injector plate, the manifold, and the catalystbeing annular and extending circumferentially around a convergentsection of the convergent-divergent nozzle.
 8. The rocket engine ofclaim 7, wherein the injector plate includes a plurality of aperturesextending therethrough.
 9. The rocket engine of claim 8, wherein theplurality of apertures are oriented parallel to a longitudinal axis ofthe rocket engine.
 10. The rocket engine of claim 9, wherein at leastsome of the plurality of apertures have aperture inlets beingcircumferentially offset from aperture outlets relative to alongitudinal axis of the rocket engine for creating a swirl in theoxidizer.
 11. The rocket engine of claim 1, wherein the injector plateis a disk and wherein the catalyst includes a mesh being disk-shaped,the mesh having at least two faces, one of the at least two faceslocated proximate the injector plate and defining the outlet of thecatalyst, the manifold extending over the other of the at least tofaces, the other of the at least two faces defining the inlet of thecatalyst.
 12. The rocket engine of claim 1, wherein the injector plateextends circumferentially all around a longitudinal axis of the rocketengine, the catalyst being annular and extending around the injectorplate and located radially outwardly of the injector plate relative to alongitudinal axis of the rocket engine, the catalyst having a meshhaving at least two faces, one of the at least two faces locatedproximate the injector plate and defining the outlet of the catalyst,the manifold extending over the other of the at least to faces, theother of the at least two faces defining the inlet of the catalyst. 13.The rocket engine of claim 12, wherein the injector plate, the manifold,and the catalyst are located at a convergent section of aconvergent-divergent nozzle of the rocket engine.
 14. An integratedinjector assembly for injecting an oxidizer within a combustion chamberof a rocket engine, comprising an inner wall defining an inlet fluidlyconnectable to a source of an oxidizer, an outer wall spaced apart fromthe inner wall to define a cavity therebetween, the outer wall definingan outlet for injecting a catalyzed oxidizer within the combustionchamber, and a catalyst located within the cavity.
 15. The integratedinjector assembly of claim 14, wherein the inner wall and the outer wallare annular and circumferentially extend all around a longitudinal axis,the inner wall located radially inwardly of the outer wall, the cavityand the catalyst being annular and extending around the longitudinalaxis.
 16. The integrated injector assembly of claim 15, wherein theinlet includes a plurality of apertures extending through the outer walland circumferentially distributed around the longitudinal axis.
 17. Theintegrated injector assembly of claim 15, wherein the outlet includes aplurality of apertures extending through the inner wall andcircumferentially distributed around the longitudinal axis.
 18. Theintegrated injector assembly of claim 17, wherein the plurality ofapertures have exit flow axes having each a circumferential componentrelative to the longitudinal axis.
 19. A method of supplying an oxidizerwithin a combustion chamber of a rocket engine, comprising: receiving anoxidizer; distributing the received oxidizer while catalyzing thereceived oxidizer; and injecting the catalyzed oxidizer within thecombustion chamber.
 20. The method of claim 19, wherein distributing thereceived oxidizer includes distributing the received oxidizer within anannular cavity extending all around a longitudinal axis of the rocketengine, and catalyzing the received oxidizer includes circulating thereceived oxidizer through a mesh extending around the longitudinal axis.